Optimized land mobile satellite configuration and steering method

ABSTRACT

A method for configuring and operating a spacecraft in an orbit that is inclined with respect to Earth&#39;s equatorial plane, the spacecraft including at least a solar array, a receive antenna, a transmit antenna, and radiator panels. The method includes the step of nominally orienting the yaw axis of the spacecraft, the roll axis of the spacecraft, and the radiator panels substantially parallel to Earth&#39;s equatorial plane, the pitch axis of the spacecraft and rotation axis of the solar array substantially parallel to Earth&#39;s polar axis, the Nadir vector in the yaw-pitch plane of the spacecraft, and the transmit antenna and receive antenna to angle φ nom . The method also includes the step of correcting the attitude of the receive antenna and the transmit antenna to maintain a desired degree of the receive antenna and the transmit antenna steered toward a coverage region on Earth&#39;s surface.

CROSS-REFERENCE TO RELATED APPLICATION

This application claims priority to U.S. Provisional Patent Application No. 60/633,962, filed Dec. 7, 2004, which is incorporated herein by reference in its entirety for all purposes.

STATEMENT AS TO RIGHTS TO INVENTIONS MADE UNDER FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

Not Applicable

FIELD OF THE INVENTION

The present invention relates generally to optimized spacecraft configurations and, more particularly, relates to the steering of spacecraft deployed in highly inclined orbits (“HIOs”) so as to provide high performance land mobile satellite services (“LMSSs”), while maintaining a favorable sun-spacecraft geometry.

BACKGROUND OF THE INVENTION

FIG. 1 depicts a typical arrangement for a conventional geostationary earth orbit (“GEO”) communications spacecraft. Normally, the yaw axis of a conventional spacecraft (such as conventional spacecraft 100) is directed at Earth's center, the yaw/roll plane is parallel to the equatorial plane, solar arrays 102 rotate about the pitch axis and orbit normal, and transponder panels 104 face in the north/south direction. Communications antennas 105 remain fixed, pointed at their Earth coverage regions.

As the angle from the sun to the spacecraft's yaw/roll plane increases, the solar radiation input to the spacecraft's heat-rejecting transponder panels increases, reducing the spacecraft's ability to reject waste heat. The increase in the sun's angle also increases the angle between the solar panels and the sun vector, thereby reducing the power produced by the solar arrays. Both of these effects potentially limit the payload that a spacecraft can carry. The typical arrangement depicted in FIG. 1 provides a conventional GEO communication spacecraft with adequate solar array power and thermal heat rejection, because it limits the angle between the sun and the spacecraft's yaw/roll plane to within the range of sun declination angles, or ±23.5°.

Despite its elegance and simplicity, a problem arises when conventional GEO spacecraft, such as spacecraft 100, are adapted to highly inclined orbits of the type presently in use to provide LMSS. For example, the SIRIUS SATELLITE RADIO® mobile satellite system provides S-band Digital Audio Radio Service (“S-DARS”) using three satellites in 24-hour period repeating ground track orbits with inclination of 63.4 degrees, eccentricity of 0.27, argument of perigee of 270°, and apogee longitude of 96° west. Using the FIG. 1 spacecraft arrangement, the maximum yaw/roll plane sun angle, or the sum of the maximum sun declination and the orbit inclination, would be 87°, resulting in insufficient power and heat rejection to operate a communications payload.

To address this problem, the SIRIUS SATELLITE RADIO® mobile satellite system satellites use a sun nadir pointing (“SNP”) yaw steering approach, rather than the conventional approach outlined above. See, e.g., R. Briskman et al., S-DARS Broadcast For Inclined elliptical Orbits, 52nd International Astronautical Congress, October 1 to 5, Toulouse, France. Using SNP, the yaw axis of the spacecraft remains pointed to Earth, while the spacecraft continuously rotates about the yaw axis to keep the sun vector in the spacecraft's yaw/roll plane. Although this strategy provides beneficial power and thermal conditions using a conventional GEO spacecraft, SNP requires that the spacecraft be capable of executing large-angle yaw rotations. FIG. 2, which is a chart depicting the required yaw rotation for SNP, illustrates that a 160° rotation is required for a sun beta angle of 10°.

Further to these disadvantages, SNP also restricts the antenna designs that can be utilized to specific types which can compensate for the spacecraft yaw rotation. For example, the SIRIUS SATELLITE RADIO® mobile satellite system S-band antenna utilizes a folded optics design that provides elliptically shaped coverage for North America. As the spacecraft rotates in yaw to follow the yaw steering profile, the antenna sub-reflector counter-rotates correspondingly, so that the elliptical pattern remains properly oriented over the coverage region.

This known rotating sub-reflector approach is not feasible for the next generation of the antennas envisioned for LMSS. The next-generation antennas may have large reflector surfaces (e.g., 12 meters compared to 2.5 meters for the conventional systems), which are contoured to provide a coverage pattern that closely matches the shape of the continental United States. The next generation approach will increase the signal directivity by up to 2 to 3 dB, increasing the quality of service, and increasing the probability that the signal strength will be sufficient for mobile receivers to operate.

Commonly-assigned U.S. Pat. No. 6,616,104 (“the Cheng patent”) describes an improved arrangement for a spacecraft which remedies some of the deficiencies of other conventional spacecraft. The Cheng patent describes a system which utilizes body steering only, fixing the antenna boresight in the spacecraft body frame, with a roll angle equal to the orbit inclination. With a fixed antenna boresight, the arrangement described in the Cheng patent is generally limited to spacecraft which operate in a narrow region about orbit apogee, such as spacecraft which operate over a limited range of latitudes and only 8 hours of a 24-hour orbit period.

Due to their substantial build, launch and operational costs, it is considered highly desirable to overcome the deficiencies of conventional LMSS spacecraft configurations and attitude steering methods. Specifically, it is desirable to provide an enhanced land mobile satellite configuration and steering method which provides a favorable sun-spacecraft geometry compatible with current and future payload power and thermal requirements, without the need for yaw steering.

SUMMARY OF THE INVENTION

The present invention relates generally to spacecraft configurations and steering and, more particularly, relates to the steering of spacecraft deployed in HIOs so as to provide high performance LMSSs, while maintaining a favorable sun-spacecraft geometry.

The optimized land mobile satellite configuration and steering method according to the present invention addresses the limitations of conventional systems by providing a spacecraft configuration, and an attitude and antenna steering approach that provides higher performance LMSS. Specifically, the enhanced spacecraft configuration and steering method provides sun-spacecraft geometry compatible with payload power and thermal requirements, without the need for yaw steering, and also provides compatibility with next-generation high-directivity LMSS transmit antenna designs.

According to one arrangement, the present invention is a method for configuring and operating a spacecraft in an orbit that is inclined with respect to Earth's equatorial plane, the spacecraft including at least a solar array, a receive antenna, a transmit antenna, and radiator panels. The method includes the step of nominally orienting the yaw axis of the spacecraft, the roll axis of the spacecraft, and the radiator panels substantially parallel to Earth's equatorial plane, the pitch axis of the spacecraft and rotation axis of the solar array substantially parallel to Earth's polar axis, the Nadir vector in the yaw-pitch plane of the spacecraft, and the transmit antenna and receive antenna to angle φ_(nom). The method also includes the step of correcting the attitude of the receive antenna and the transmit antenna to maintain a desired degree of the receive antenna and the transmit antenna steered toward a coverage region on Earth's surface. The correcting step further includes the steps of offsetting the attitude of the spacecraft by angle θ with respect to the pitch axis of the nominal orientation, and by angle φ_(b) with respect to the roll axis of the spacecraft, and offsetting the receive antenna and the transmit antenna at angle φ_(a) with respect to the spacecraft.

According to a second arrangement, the present invention is a spacecraft operable in an inclined orbit with respect to Earth's equator, including at least one solar array rotatable about the pitch axis of the spacecraft, a receive antenna, and a transmit antenna further including an antenna attitude adjuster. The spacecraft also includes at least one radiator panel, and a spacecraft attitude adjustor. The spacecraft is configured such that the spacecraft has a nominal orientation in which the yaw axis of the spacecraft, the roll axis of the spacecraft, and the radiator panels are substantially parallel to Earth's equatorial plane, in which the pitch axis of the spacecraft is substantially parallel to Earth's polar axis, in which the Nadir vector is in the yaw-pitch plane of the spacecraft, and in which the transmit antenna and receive antenna are oriented at angle φ_(nom). The antenna attitude adjustor and the spacecraft attitude adjustor correct the attitude of the transmit antenna to maintain a desired degree of the receive antenna and the transmit antenna steered toward a coverage region on Earth's surface, by offsetting the attitude of spacecraft by angle θ with respect to the pitch axis of the nominal orientation, and by angle φ_(b) with respect to the roll axis of the spacecraft, and by offsetting the receive antenna and the transmit antenna at angle φ_(a) with respect to the spacecraft.

Angle φ_(nom) is based upon orbit parameters, orbital region of broadcast service, and/or the coverage region. Offset angles are determined based upon maximum allowable sun angle, or roll offset from the nominal orientation. The coverage region is North America, although in alternate embodiments of the present invention the coverage region is another region aside from North America.

In contrast to conventional designs, the optimized spacecraft configuration and steering method according to the present invention applies to spacecraft that may operate for extended periods over a wide range of latitudes, such as 16 of 24 hours, and over the entire Northern hemisphere, as required for the SIRIUS SATELLITE RADIO® mobile satellite system. The approach used by the present invention partitions the steering to both the spacecraft and the antenna to provide improved sun geometry for power generation and thermal control. In doing so, the present invention represents a significant improvement over the conventional methods.

In the following description of the preferred embodiment, reference is made to the accompanying drawings that form a part thereof, and in which is shown by way of illustration a specific embodiment in which the invention may be practiced. It is to be understood that other embodiments may be utilized and changes may be made without departing from the scope of the present invention.

BRIEF DESCRIPTION OF THE DRAWINGS

Referring now to the drawings in which like reference numbers represent corresponding parts throughout:

FIG. 1 depicts the a typical arrangement for a conventional geostationary earth orbit (“GEO”) communications spacecraft;

FIG. 2 depicts the required yaw rotation for SNP;

FIG. 3 is a diagram depicting the optimized spacecraft configuration according to one embodiment of the present invention;

FIGS. 4A and 4B illustrate the method for configuring and operating a spacecraft in an orbit which is inclined with respect to Earth's equatorial plane;

FIG. 5 illustrates attitude steering according to the present invention, showing the spacecraft in-orbit geometry in a typical highly inclined orbit where LMSS is provided to a coverage region in the Northern hemisphere;

FIGS. 6A and 6B depict the total roll offset, the antenna roll offset, the body roll offset, and the body pitch offset for a conventional orbit;

FIGS. 7A and 7B depict the total roll offset, the antenna roll offset, the body roll offset, and the body pitch offset for an “Option 1” orbit;

FIGS. 8A and 8B depict the total roll offset, the antenna roll offset, the body roll offset, and the body pitch offset for an “Option 2” orbit;

FIGS. 9A and 9B depict the total roll offset, the antenna roll offset, the body roll offset, and the body pitch offset for an “Option 3” orbit; and

FIG. 10 is a block diagram of a system that implements the attitude and antenna steering approach according to the present invention.

DETAILED DESCRIPTION OF THE INVENTION

The present invention provides an LMSS for North American coverage, which overcomes the deficiencies of conventional MSS systems. Specifically, the present invention provides an enhanced LMSS using an optimized land mobile satellite configuration and steering method, which provides a favorable sun-spacecraft geometry compatible with current and future payload power and thermal requirements, without the need for yaw steering.

FIG. 3 is a diagram depicting the optimized spacecraft configuration according to one embodiment of the present invention. Briefly, the present invention is a method for configuring and operating a spacecraft in an orbit that is inclined with respect to Earth's equatorial plane, the spacecraft including at least a solar array, a receive antenna, a transmit antenna, and radiator panels. The method includes the step of nominally orienting the yaw axis of the spacecraft, the roll axis of the spacecraft, and the radiator panels substantially parallel to Earth's equatorial plane, the pitch axis of the spacecraft and rotation axis of the solar array substantially parallel to Earth's polar axis, the Nadir vector in the yaw-pitch plane of the spacecraft, and the transmit antenna and receive antenna to angle φ_(nom). The method also includes the step of correcting the attitude of the receive antenna and the transmit antenna to maintain a desired degree of the receive antenna and the transmit antenna steered toward a coverage region on Earth's surface. The correcting step further includes the steps of offsetting the attitude of the spacecraft by angle θ with respect to the pitch axis of the nominal orientation, and by angle φ_(b) with respect to the roll axis of the spacecraft, and offsetting the receive antenna and the transmit antenna at angle φ_(a) with respect to the spacecraft.

In more detail, spacecraft 301 includes solar arrays 302 and 303, receive antenna 305, transmit antenna 306, and radiator panels 307. Inherent to spacecraft 301 are yaw axis X_(b), roll axis Y_(b), pitch axis Z_(b), and Nadir vector N (not shown).

FIGS. 4A and 4B are a flow charts illustrating the method for configuring and operating spacecraft 301 in an orbit which is inclined with respect to Earth's equatorial plane. As depicted in FIG. 4A, the process begins (step S401), and yaw axis X_(b) of spacecraft 301, roll axis Y_(b) of spacecraft 301, and radiator panels 307 are nominally oriented substantially parallel to Earth's equatorial plane, and pitch axis Z_(b) of spacecraft 301 and the rotation axis of solar arrays 302 and 303 are nominally oriented substantially parallel to Earth's polar axis (step S402). Additionally, the Nadir vector N is oriented in the yaw-pitch plane of spacecraft 301, and transmit antenna 306 and receive antenna 307 are oriented to angle φ_(nom) (step S402).

The attitude of receive antenna 305 and transmit antenna 306 are corrected to maintain a desired degree of receive antenna 305 and transmit antenna 306 steered toward a coverage region on Earth's surface (step S403), and the process ends (step S408).

As shown in FIG. 4B, the correcting step (step S403) begins at sub-step S404, and further includes the sub-steps of offsetting the attitude of spacecraft 301 by angle θ with respect to the pitch axis of the nominal orientation, and by angle φ_(b) with respect to the roll axis of the spacecraft, (sub-step S405), and offsetting receive antenna 305 and transmit antenna 306 at angle φ_(a) with respect to the spacecraft (sub-step S406). Sub-step S407 ends correcting step S403.

Offset angles are determined based upon maximum allowable sun angle, or roll offset from the nominal orientation. The coverage region is North America, although in alternate aspects of the present invention, the coverage region includes areas outside of North America.

In contrast to conventional designs, the optimized spacecraft configuration and steering method according to the present invention applies to spacecraft that may operate for extended periods over a wide range of latitudes, such as 16 of 24 hours, and over the entire Northern hemisphere, as required for the SIRIUS SATELLITE RADIO® mobile satellite system. The approach used by the present invention partitions the steering to both the spacecraft and the antenna to provide improved sun geometry for power generation and thermal control. In doing so, the present invention represents a significant improvement over the conventional methods.

As shown in FIG. 3, spacecraft 301 includes both receive antenna 305 and transmit antenna 306 that are oriented with their boresights canted at a specified angle φ_(nom) about the body roll (Y_(b)) axis. The antenna design facilitates a roll rotation Δφ_(scan) over a specified range about the nominal roll offset. The nominal angle φ_(nom) is selected based on the orbit parameters, the orbital region of broadcast service, and the coverage region to which service is to be provided.

FIG. 5 further illustrates attitude steering according to the present invention, showing the spacecraft in-orbit geometry in a typical highly inclined orbit where LMSS is provided to a coverage region in the Northern hemisphere. As shown in FIG. 5, to provide antenna pointing to the coverage region, the antenna boresight tracks a commanded attitude represented by the 3×3 transformation matrix expressed below in Equation (1): T _(c) =T _(φ) T _(θ) T _(ENP)  (1)

In Equation (1), T_(φ) is the roll correction matrix, T_(θ) is the pitch correction matrix, and T_(ENP) is the Equatorial Nadir Pointing (“ENP”) matrix. Antenna pointing is accomplished by a combination of spacecraft body steering and antenna boresight steering, as described more fully below.

The matrix T_(ENP) describes the nominal spacecraft body pointing direction in inertial space in the absence of additional body pointing corrections, and is expressed below in Equation (2): T _(ENP) =[X _(i) Y _(i) Z _(i)]  (2)

In Equation (2), Z_(i) is the inertial frame earth pole axis vector, and vectors X_(i) and Y_(i) are expressed by Equation (3), below: $\begin{matrix} {{Y_{i} = \frac{Z_{i} \times N}{{Z_{i} \times N}}}{X_{i} = \frac{Y_{i} \times Z_{i}}{{Y_{i} \times Z_{i}}}}} & (3) \end{matrix}$

In Equation (3), N is the inertial frame spacecraft Nadir vector. Aligning the body frame with the inertial attitude represented by T_(ENP) causes the spacecraft yaw/roll plane to remain parallel to Earth equatorial plane and the spacecraft pitch axis to remain aligned with Earth pole axis. The Nadir vector remains in the spacecraft yaw/pitch plane, and when the spacecraft crosses the equator, the Nadir vector is aligned with the spacecraft yaw axis. As can be seen from FIG. 4, the ENP pointing strategy results in sun angles within ±23.5°, which are compatible with existing GEO spacecraft designs.

Starting from the ENP attitude, as the spacecraft orbital position changes, antenna pointing is accomplished by commanding the spacecraft body and antenna pointing offsets represented by the matrices of Equation (4): $\begin{matrix} {T_{\phi} = {{\begin{bmatrix} {\cos\quad\phi} & {0\quad} & {{- \sin}\quad\phi} \\ 0 & 1 & 0 \\ {\sin\quad\phi} & \quad & {\cos\quad\phi} \end{bmatrix}\quad T_{\theta}} = \begin{bmatrix} {\cos\quad\theta} & {\sin\quad\theta} & 0 \\ {{- \sin}\quad\theta} & {\cos\quad\theta} & 0 \\ 0 & 0 & 1 \end{bmatrix}}} & (4) \end{matrix}$

In Equation (4), pitch and roll angles θ, φ are as expressed below in Equation (5): $\begin{matrix} {{\theta = {\tan^{- 1}\left( \frac{Y_{i} \cdot V_{t}}{X_{i} \cdot V_{t}} \right)}}{\phi = {- {\sin^{- 1}\left( {Z_{i} \cdot V_{t}} \right)}}}} & (5) \end{matrix}$

In Equation (5), V_(t) is a unit vector from the spacecraft to a specified location on Earth's surface in the coverage region, such that the antenna remains correctly pointed. V_(t), for example, could be the sub-satellite point when the spacecraft is at orbit apogee. The pitch correction is generally implemented as a spacecraft body rotation from the ENP attitude, and does not impact the sun angles from the spacecraft yaw/roll plane. The roll rotation is generally implemented as a combination of body and antenna pointing corrections, as expressed in Equation (6): φ=φ_(b)φ_(a)

The body and antenna roll corrections φ_(b) and φ_(a) may be determined in any number of ways, depending on the maximum allowable antenna roll scan angle range or the allowable sun angles from the spacecraft yaw/roll plane. One method for partitioning the offsets based on the maximum antenna scan angle Δφ_(scan) is φ_(a) =c(φ−φ_(nom))+φ_(nom) φ_(b)=(1−c)(φ−φ−_(nom))  (7)

In Equation (7), c and φ_(nom) are expressed as shown below in Equation (8): $\begin{matrix} {{c = \frac{\Delta\quad\phi_{scan}}{\Delta\quad\phi}}{{\Delta\quad\phi} = \frac{\phi_{\max} - \phi_{\min}}{2}}{\phi_{nom} = \frac{\phi_{\max} + \phi_{\min}}{2}}} & (8) \end{matrix}$

In Equations (7) and (8), φ_(nom) is the nominal antenna roll offset in the spacecraft body frame, and Δφ is the roll angle excursion required about the nominal offset, i.e., φ_(nom)−Δφ≦φ_(a)≦φ_(om)+Δφ. These quantities are computed based on φ_(max) and φ_(min), which are the maximum and minimum roll offsets required from the ENP frame over the broadcast service region of the orbit (e.g., the Northern hemisphere).

Using the pointing strategy outlined above, the maximum spacecraft body roll offset is (1−c)Δφ, and the maximum sun angle from the spacecraft yaw/roll plane is the sum of this angle and the sun declination of 23.5°. Alternate methods for partitioning the antenna and body roll pointing offsets are also contemplated by the present invention, such as partitioning based on the maximum allowable sun angle or roll offset from the ENP frame.

Examples of the steering profiles are now given for the orbits given in Table 1, which includes the Sirius Satellite Radio orbit disclosed in U.S. Pat. No. 6,564,053 (“the Briskman patent”). Several other optimized orbits which are selected to provide enhanced elevation angles and probability of signal availability for LMSS, denoted as “Option 1,” “Option 2,” and “Option 3”, are also listed.

The optimized orbits listed in Table 1 have substantially teardrop-shaped or oval-shaped ground tracks, and are optimized based upon-performance criteria. A full explanation of the methodology and performance criteria used to select the optimized orbits is available in U.S. patent application Ser. No. 11/030,282, entitled “Optimized Land Mobile Satellite System For North American Coverage,” filed Jan. 7, 2005 (“the Goodzeit '282 application”). Both the Briskman patent and the Goodzeit '282 application are incorporated by reference herein in the entirety for all purposes, although further details relating to the content of both references are omitted for the sake of brevity. TABLE 1 Orbits for North American LMSS Parameter SIRIUS Orbit Option 1 Option 2 Option 3 Inclination 63.4 48.2 45.4 58.9 Eccentricity 0.27 0.24 0.40 0.35 Apogee 96 95 95 95 longitude (° W) Antenna roll 27.5 19.2 25.5 28.7 offset (°) φ_(nom) Roll angle 35.9 29.0 19.8 30.2 variation about nominal (°) Δφ Maximum body 20.9 14.0 4.8 15.2 roll offset (°) Maximum pitch 2.2 3.2 4.8 3.0 angle (°)

All of the orbits listed in Table 1 have a semi-major axis of 42,164 km and an argument of perigee of 270°. Also provided in Table 1 are the nominal antenna roll offset, the roll angle variation about the nominal offset, and the maximum body roll offset, based on Equations (7) and (8). The results of these equations assume an antenna roll scan angle range of ±15°, where the spacecraft operates for ±8 hours about orbit apogee, and where the vector V_(t) in Equation (5) is from the spacecraft to the sub-satellite point at orbit apogee. Table 1 also provides the maximum pitch angle, which is computed using Equation (5).

FIG. 6A is a chart depicting the total roll offset (denoted by reference 601), the antenna roll offset (denoted by reference 602), and the body roll offset (denoted by reference 603) for the orbit described in the Briskman patent, and FIG. 6B is a chart depicting the body pitch offset (denoted by reference 604) for the same orbit.

FIG. 7A is a chart depicting the total roll offset (denoted by reference 701), the antenna roll offset (denoted by reference 702), and the body roll offset (denoted by reference 703) for the Option 1 orbit, and FIG. 7B is a chart depicting the body pitch offset (denoted by reference 704) for the same orbit.

FIG. 8A is a chart depicting the total roll offset (denoted by reference 801), the antenna roll offset (denoted by reference 802), and the body roll offset (denoted by reference 803) for the Option 2 orbit, and FIG. 8B is a chart depicting the body pitch offset (denoted by reference 804) for the same orbit.

FIG. 9A is a chart depicting the total roll offset (denoted by reference 901), the antenna roll offset (denoted by reference 902), and the body roll offset (denoted by reference 903) for the Option 3 orbit, and FIG. 9B is a chart depicting the body pitch offset (denoted by reference 904) for the same orbit.

As can be seen from Table 1 and FIGS. 6 to 9, the maximum required roll and pitch offsets of 20.9° and 4.8° necessary to implement the steering approach according to the present invention are far smaller than the yaw offsets required to implement a conventional SNP steering approach, which is shown in FIG. 2. The reduced spacecraft maneuvering requirements improve attitude pointing performance and allow the use of smaller and lighter reaction wheels assemblies (“RWAs”). Also, it should be noted that the optimized orbits result in a smaller body roll offsets than the conventional orbit, thereby reducing the sun angle from the yaw/roll plane and improving spacecraft power and thermal performance.

FIG. 10 is a block diagram of a system that implements the attitude and antenna steering approach according to the present invention. Inertial Attitude Determination Logic 1001 computes the spacecraft inertial attitude using gyro angular rate data from Inertial Measurement Unit (“IMU”) 1002, and attitude data from star trackers 1004 and sun sensors 1005. Those skilled in the art will recognize that other combinations of attitude sensors may be used for attitude determination including earth sensors and other sun sensors.

Spacecraft Ephemeris Generator 1006 computes the spacecraft orbital position and velocity and provides an estimate of the Nadir vector N, the vector from the spacecraft to Earth's center. Using this vector, the Attitude And Antenna Steering Logic 1007 computes the commanded spacecraft inertial-to-body frame transformation and commanded antenna roll angle according to the steps described above in Equations (1) to (8).

As described in Equation (6), a portion of the roll pointing command is allocated to the antenna and a portion is allocated to the spacecraft body. In consideration of this partitioning, Equation (1) can also be expressed as Equation (9): T _(c) =T _(φ) _(a) T _(φ) _(b) T _(θ) T _(ENP)  (9)

The spacecraft body portion of the commanded attitude T_(θ) _(b) T_(θ)T_(ENP) shown in FIG. 9, is input to the Attitude Error Logic 1008, which computes the attitude pointing and rate errors representative of the error between the present spacecraft attitude and the commanded inertial attitude. The attitude and rate errors are input to a Proportional-Integral-Derivative (“PID”) controller 1009 that computes torque commands for attitude control. The torque commands are then input to RWA Torque Distribution Logic 1010 that commands individual RWAs 1011 to produce control torque that causes the spacecraft to track the commanded attitude.

The antenna roll pointing correction represented by T_(φ) _(a) or φ_(a) is input to the Antenna Pointing Control Logic 1012. This logic commands the Antenna Pointing Electronics 1013 to achieve the commanded antenna boresight roll pointing. The antenna pointing is controlled either mechanically or electrically, depending on the antenna design. For example, the antenna may include a gimbaled reflector that may be rotated to change the antenna boresight pointing. Alternatively, the antenna may be a phased array antenna design, such that the phases of the individual antenna phase elements are commanded to achieve the desired pointing.

The steering method according to the invention may be augmented with additional spacecraft body and antenna pointing corrections to maximize LMSS signal strength to the overage region as the spacecraft position in orbit changes. Furthermore, the coverage pattern can be enlarged or decreased to compensate for the effects of changes in spacecraft altitude. Additionally, although the steering of antennas has been described above using mechanical steering, other methods, such as electrical steering using a phased array antenna, are also contemplated by the present invention. All of these enhancements are possible without departing from the spirit and scope of the invention.

The invention has been described with particular illustrative embodiments. It is to be understood that the invention is not limited to the above-described embodiments and that various changes and modifications may be made by those of ordinary skill in the art without departing from the spirit and scope of the invention. 

1. A method for configuring and operating a spacecraft in an orbit that is inclined with respect to Earth's equatorial plane, the spacecraft including at least a solar array, a receive antenna, a transmit antenna, and radiator panels, the method comprising the steps of: nominally orienting the yaw axis of the spacecraft, the roll axis of the spacecraft, and the radiator panels substantially parallel to Earth's equatorial plane, the pitch axis of the spacecraft and rotation axis of the solar array substantially parallel to Earth's polar axis, the Nadir vector in the yaw-pitch plane of the spacecraft, and the transmit antenna and receive antenna to angle φ_(nom); and correcting the attitude of the receive antenna and the transmit antenna to maintain a desired degree of the receive antenna and the transmit antenna steered toward a coverage region on Earth's surface, further comprising the steps of: offsetting the attitude of the spacecraft by angle θ with respect to the pitch axis of the nominal orientation and by angle φ_(b) with respect to the roll axis of the spacecraft; and offsetting the receive antenna and the transmit antenna at angle φ_(a) with respect to the spacecraft.
 2. The method according to claim 1, wherein angle φ_(nom) is based upon orbit parameters, orbital region of broadcast service, and/or the coverage region.
 3. The method according to claim 1, wherein in the correcting step, the receive antenna and the transmit antenna track a commanded attitude expressed by: T _(c) =T _(φ) T _(θ) T _(ENP), wherein T_(φ) is a roll correction matrix, T_(θ) is a pitch correction matrix, and T_(ENP) is expressed by: T _(ENP) =[X _(i) Y _(i) Z _(i)], wherein Z_(i) is the Earth's polar axis vector, and vectors X_(i) and Y_(i) are expressed by: $Y_{i} = \frac{Z_{i} \times N}{{Z_{i} \times N}}$ ${X_{i} = \frac{Y_{i} \times Z_{i}}{{Y_{i} \times Z_{i}}}},$ wherein N is a vector from the spacecraft to Earth's center, and wherein T_(φ) and T_(θ) are expressed by: ${T_{\phi} = {{\begin{bmatrix} {\cos\quad\phi} & {0\quad} & {{- \sin}\quad\phi} \\ 0 & 1 & 0 \\ {\sin\quad\phi} & \quad & {\cos\quad\phi} \end{bmatrix}\quad T_{\theta}} = \begin{bmatrix} {\cos\quad\theta} & {\sin\quad\theta} & 0 \\ {{- \sin}\quad\theta} & {\cos\quad\theta} & 0 \\ 0 & 0 & 1 \end{bmatrix}}},$ wherein angles θ and φ are expressed by: $\theta = {\tan^{- 1}\left( \frac{Y_{i} \cdot V_{t}}{X_{i} \cdot V_{t}} \right)}$ ϕ = −sin⁻¹(Z_(i) ⋅ V_(t)), and wherein V_(i) is a unit vector from the spacecraft to a location within the coverage region.
 4. The method according to claim 1, wherein angle φ is expressed by: φ=φ_(b)+φ_(a), wherein φ_(b) and φ_(a) are expressed by: φ_(a) =c(φ−φ_(nom))+φ_(nom) φ_(b)=(1−c)(φ−φ_(nom)), wherein c and φ_(nom) is expressed by: $c = \frac{\Delta\quad\phi_{scan}}{\Delta\quad\phi}$ ${\Delta\quad\phi} = \frac{\phi_{\max} - \phi_{\min}}{2}$ ${\phi_{nom} = \frac{\phi_{\max} + \phi_{\min}}{2}},$ wherein Δφ_(scan) is maximum antenna scan angle, φ_(max) is maximum roll offset from the nominal orientation over at least a portion of the orbit, and φ_(min) is minimum roll offset from the nominal orientation over at least a portion of the orbit.
 5. The method according to claim 1, wherein angle φ_(b) and/or angle φ_(a) are determined based upon maximum allowable sun angle, or roll offset from the nominal orientation.
 6. The method according to claim 4, wherein the orbit has a semi-major axis of 42,164 km and an argument of perigee of 270°, and wherein φ_(nom) is 19.2°, φ_(scan) is 29.0°, φ_(max) is 14.0°, and φ_(min) is 3.2°.
 7. The method according to claim 4, wherein the orbit has a semi-major axis of 42,164 km and an argument of perigee of 270°, and wherein φ_(nom) is 25.5°, φ_(scan) is 19.8°, φ_(max) is 4.8°, and φ_(min) is 4.8°
 8. The method according to claim 4, wherein the orbit has a semi-major axis of 42,164 km and an argument of perigee of 270°, and wherein φ_(nom) is 28.7°, φ_(scan) is 30.2°, φ_(max) is 15.2°, and φ_(min) is 3.0°
 9. The method according to claim 1, wherein the coverage region is North America.
 10. A spacecraft operable in an inclined orbit with respect to Earth's equator, comprising: at least one solar array rotatable about the pitch axis of the spacecraft; a receive antenna; a transmit antenna further comprising an antenna attitude adjuster; at least one radiator panel; and a spacecraft attitude adjustor, wherein the spacecraft is configured such that the spacecraft has a nominal orientation in which the yaw axis of the spacecraft, the roll axis of the spacecraft, and the radiator panels are substantially parallel to Earth's equatorial plane, in which the pitch axis of the spacecraft is substantially parallel to Earth's polar axis, in which the Nadir vector is in the yaw-pitch plane of the spacecraft, and in which the transmit antenna and receive antenna are oriented at angle φ_(nom), and wherein the antenna attitude adjustor and the spacecraft attitude adjustor correct the attitude of the transmit antenna to maintain a desired degree of the receive antenna and the transmit antenna steered toward a coverage region on Earth's surface, by offsetting the attitude of spacecraft by angle θ with respect to the pitch axis of the nominal orientation, and by angle φ_(b) with respect to the roll axis of the spacecraft, and by offsetting the receive antenna and the transmit antenna at angle φ_(a) with respect to the spacecraft.
 11. The spacecraft according to claim 10, wherein angle φ_(nom) is based upon orbit parameters, orbital region of broadcast service, and/or the coverage region.
 12. The spacecraft according to claim 10, wherein in the correcting step, the receive antenna and the transmit antenna track a commanded attitude expressed by: T _(c) =T _(φ) T _(θ) T _(ENP), wherein T_(φ) is a roll correction matrix, T_(θ) is a pitch correction matrix, and T_(ENP) is expressed by: T _(ENP) =[X _(i) Y _(i) Z _(i)], wherein Z_(i) is the Earth's polar axis vector, and vectors X_(i) and Y_(i) are expressed by: $Y_{i} = \frac{Z_{i} \times N}{{Z_{i} \times N}}$ ${X_{i} = \frac{Y_{i} \times Z_{i}}{{Y_{i} \times Z_{i}}}},$ wherein N is a vector from the spacecraft to Earth's center, and wherein T_(φ) and T_(θ) are expressed by: ${T_{\phi} = {{\begin{bmatrix} {\cos\quad\phi} & {0\quad} & {{- \sin}\quad\phi} \\ 0 & 1 & 0 \\ {\sin\quad\phi} & \quad & {\cos\quad\phi} \end{bmatrix}\quad T_{\theta}} = \begin{bmatrix} {\cos\quad\theta} & {\sin\quad\theta} & 0 \\ {{- \sin}\quad\theta} & {\cos\quad\theta} & 0 \\ 0 & 0 & 1 \end{bmatrix}}},$ wherein angles θ and φ are expressed by: $\theta = {\tan^{- 1}\left( \frac{Y_{i} \cdot V_{t}}{X_{i} \cdot V_{t}} \right)}$ ϕ = −sin⁻¹(Z_(i) ⋅ V_(t)), and wherein V_(i) is a unit vector from the spacecraft to a location within the coverage region.
 13. The spacecraft according to claim 10, wherein angle φ is expressed by: φ=φ_(b)+φ_(a), wherein φ_(b) and φ_(a) are expressed by: φ_(a) =c(φ−φ_(nom))+φ_(nom) φ_(b)=(1−c)(φ−φ_(nom)), wherein c and φ_(nom) is expressed by: $c = \frac{{\Delta\phi}_{scan}}{\Delta\phi}$ ${{\Delta\phi} = \frac{\phi_{\max} - \phi_{\min}}{2}},{\phi_{nom} = \frac{\phi_{\max} + \phi_{\min}}{2}}$ wherein Δφ_(scan) is maximum antenna scan angle, φ_(max) is maximum roll offset from the nominal orientation over at least a portion of the orbit, and φ_(min) is minimum roll offset from the nominal orientation over at least a portion of the orbit.
 14. The spacecraft according to claim 10, wherein angle φ_(b) and/or angle φ_(a) are determined based upon maximum allowable sun angle, or roll offset from the nominal orientation.
 15. The spacecraft according to claim 13, wherein the orbit has a semi-major axis of 42,164 km and an argument of perigee of 270°, and wherein φ_(nom) is 19.2°, φ_(scan) is 29.0°, φ_(max) is 14.0°, and φ_(min) is 3.2°.
 16. The spacecraft according to claim 13, wherein the orbit has a semi-major axis of 42,164 km and an argument of perigee of 270°, and wherein φ_(nom) is 25.5°, φ_(scan) is 19.8°, φ_(max) is 4.8°, and φ_(min) is 4.8°
 17. The spacecraft according to claim 13, wherein the orbit has a semi-major axis of 42,164 km and an argument of perigee of 270°, and wherein φ_(nom) is 28.7°, φ_(scan) is 30.2°, φ_(max) is 15.2°, and θ_(min) is 3.0° 